Aircraft engines of this type are known. Aircraft engines with a dual-shaft design have two shafts seated to rotate coaxially and independently of one another, and each has respectively a separate compressor and a separate turbine. The combustion chamber is disposed between the medium-pressure compressor and the medium-pressure turbine, and in turbojet engines, the propulsion jet is disposed downstream of the medium-pressure turbine. In contrast, in aircraft engines with a propeller the propulsive power is essentially generated by means of the transmission of a shaft output to a propeller operating in atmospheric air.
The embodiment of the by-pass flow of dual-flow turbojet engines typically includes the turbofan operating in the by-pass flow and whose blades can either form an extension of the compressor blades (front fan) or the turbine blades (aft fan) of the main rotor, or is driven by a separate rotor. The air flow entering into the engine is guided in a hot primary flow and in a second, cold secondary flow that is essentially independent of the primary flow, and the primary flow is surrounded coaxially and at least partly along its length by the secondary flow in annular or sheath-type fashion. In a bypass engine with a dual-shaft embodiment, this secondary flow is branched off from the primary air flow in a controlled manner and mixed again with the actual propulsion stream in the nozzle.
The primary flow is compressed in the low-pressure compressor and subsequently in the medium-pressure compressor. Afterward, heat energy is supplied to it in the combustion chamber by means of the combustion of continuously injected fuel with the oxygen of the compressed air. Only as much energy is drawn from the compressed and heated air or combustion gas by the downstream turbines as is required to drive the compressors and the turbofan or the propeller.
In comparison to conventional turbojet engines, dual-flow turbojet engines of this type have an increased propulsion effect, an improved ability to be controlled and a reduced specific fuel consumption, particularly at average flying speeds. Thus, any further improvement in the performance of aircraft engines of this type no longer appears to be possible.
Stationary gas turbine systems are known in which a dynamic pressure machine operating with isochoric combustion is disposed between the compressor and the gas turbine. This dynamic pressure machine has a cellular wheel that rotates between respectively a side part on the air side and a side part on the gas side, both provided with input and output openings, and has a number of cells in which a continuously repeating ignition and combustion process takes place. If, during operation, the dynamic pressure machine produces a homogenous fuel distribution in the longitudinal direction of the cells (European Patent Disclosure EP 0468083), problems occur when a low-pressure turbine and a high-pressure turbine are connected to this dynamic pressure machine, because both turbines are to be operated at the same optimum temperature if possible, although the gases are supplied to the high- and low-pressure turbines at different pressures. Therefore a change has been made to generate two partial quantities of gas enriched to different extents with fuel from compressed mass flows that are released from the cells after ignition and compression, wherein the partial quantity enriched to a lesser extent with fuel is supplied to the high-pressure turbine, and the second partial quantity enriched to a greater extent with fuel is supplied to the low-pressure turbine. It has been achieved by means of the different degrees of enrichment of the partial quantities with fuel, that neither overly hot gases are supplied to the high-pressure turbine, nor are overly cold gases supplied to the low-pressure turbine. This results in an improvement in the efficiency of the gas turbine systems.